In general, the invention relates to high temperature components that are covered by protective coatings, and are cooled by various air-flow systems. In some specific embodiments, the high temperature components are part of a gas turbine engine.
Turbine systems are widely utilized in fields such as power generation. A conventional gas turbine system utilized for power generation includes a compressor, a combustor, and a turbine. Typically, such a gas turbine system produces high temperature flows of gas through a flow path defined by the components of the turbine. Higher temperature flows generally are desirable, as they can lead to increased performance, efficiency, and power output of the gas turbine system. The high temperature flows are typically associated with or indicative of the types of combustion and flow conditions associated with a properly functioning gas turbine system. (In general, during gas turbine operation, for example, combustion gases may exceed about 1,600-1,700° C.; which is higher than the melting points of the engine components).
As might be expected, such high temperatures can cause excessive heating of the components within the flow path. Such heating may in turn cause one or more of these components to become damaged or to move outside of “specification”, leading to a shortened operational life. Thus, because of the desirability of these high temperature flow conditions in a properly running system, the components that are subjected to high temperature flows must be cooled to allow the gas turbine system to operate with flows at increased temperatures.
A number of strategies may be employed for cooling components that are subjected to high temperature flows. These components are typically known as “hot gas path components”. However, many of the cooling strategies employed result in comparatively low heat transfer rates and non-uniform component temperature profiles, which may be insufficient to achieve the desired cooling. Some of the cooling strategies may also decrease the overall turbine efficiency, because they divert an excessive amount of cooling air from the compressor of the engine.
For additional protection from the high-temperature gas flow, the exposed outer walls of the hot gas path components may be covered with a thermal barrier coating (TBC) system, which provides thermal insulation. TBC systems usually include at least one ceramic overcoat, and an underlying metallic bond coat. The benefits of thermal barrier coating systems are well-known.
In most of these exemplary gas turbine engine components, thin walls of high strength superalloy metals are typically used for enhanced durability, while minimizing the need for cooling thereof. Various cooling circuits and features are tailored for these individual components in their corresponding environments in the engine. For example, a series of internal cooling passages, or serpentines, may be formed in a hot gas path component. A cooling fluid may be provided to the serpentines from a plenum, and the cooling fluid may flow through the passages, cooling the hot gas path component substrate and coatings. However, this cooling strategy can sometimes result in comparatively low heat transfer rates, and non-uniform component temperature profiles.
Micro-channel cooling has the potential to significantly reduce cooling requirements, by placing the cooling features as close as possible to the heat zone. In this manner, the temperature delta between the “hot side” and “cold side” of the main load-bearing substrate material of a component can be considerably reduced, for a given heat transfer rate. The formation and use of micro cooling channels is described in a pending U.S. application Ser. No. 12/953,177 (Ronald Bunker et al), filed on Nov. 23, 2010, and assigned to the assignee of the present Application. Additional details regarding these channels are provided below. In general, the channels are formed in an external surface of the hot gas path component, and are designed to allow the passage of a cooling fluid, such as compressed air, originating in the engine compressor. The flow of the cooling fluid may thereby cool adjacent or proximate regions of the component, through convective cooling. As an example, this type of cooling system can transfer heat from the component, or from one or more of the protective layers disposed on the component, to the cooling medium.
While the use of microchannels can provide the attributes presented above, some drawbacks remain in this type of cooling system scheme—especially in the case of gas turbine engine components. As an example, in some instances, the deposition of protective layers over the channels usually requires the use of a sacrificial material to fill the channels and underlying passage holes, prior to the deposition process. The necessary removal of the sacrificial material, e.g., by leaching, after the coatings have been applied, can be a slow process. There are a limited number of outlets for the sacrificial material, like the lower access sites for the passage holes; and these outlets are relatively small.
Moreover, in this type of cooling system, the TBC system is especially important, for protecting the substrate from adverse environmental and thermal effects. (The TBC also provides an aerodynamically smooth surface for coolant flow). However, the loss of portions of the TBC system—by damage or general coating failure—will leave the underlying micro-channel exposed on its outside surface, and thereby subject to direct exposure to the hot gas temperatures. This in turn can lead to serious damage of the component.
With these considerations in mind, new methods and structures for improving cooling capabilities in gas turbine engines and other high temperature components would be welcome in the art. The innovations should enhance the performance of the cooling stream, using microchannels and cooling passage holes, and without significantly decreasing engine efficiency. Moreover, there is considerable interest in improving manufacturing processes used in the formation of the cooling system and protective coating systems. Furthermore, cooling system structures that would provide additional coolant flow in the event of partial TBC failure would also be of considerable value. The film cooling structures should also not interfere with the strength and integrity of the turbine engine part.